Abstract: A rocket propulsion system comprising a first cryogenic tank and a second cryogenic tank, wherein the first cryogenic tank is filled with a first propellant, and the second cryogenic tank is filled with a second propellant, for purposes of feeding at least one repeatedly ignitable main propulsion unit in a propulsion phase of the rocket propulsion system. For purposes of tank pressurization via at least a low level of acceleration in a ballistic phase, a first auxiliary propulsion unit can be operated by means of a first gas pressure accumulator, and at least one further auxiliary propulsion unit can be operated by means of a further gas pressure accumulator, and the rocket propulsion system is assigned an energy conversion unit, which is designed at least to charge the first and the second gas pressure accumulator, preferably in the ballistic phase.

Abstract: A solid propellant propulsion motor may comprise: a forward propellant grain extending along a longitudinal axis of a motor case between a forward end of the motor case and a first burn inhibitor layer in the motor case; the first burn inhibitor layer disposed axially adjacent to the forward propellant grain; an aft propellant grain disposed axially adjacent to the first burn inhibitor layer; a second burn inhibitor layer disposed axially adjacent to an aft end of the aft propellant grain; and an ablative material layer disposed on a radially inner surface of the aft propellant grain.

Abstract: A rocket motor for a gun-fired projectile is configured stiffen the burnable propellant in the rocket motor during burning and/or protect the rocket motor from the pressure that occurs during firing of the projectile from the gun. The rocket motor may include a rigid structure that is integrated into the burnable propellant grain to stabilize the burnable propellant grain during burning of the burnable propellant grain. The rigid structure has a matrix or truss-like shape that extends into the depth of the burnable propellant grain. The rocket motor may include a baffled end cap that covers a nozzle of the rocket motor. The end cap defines a baffled path through the end cap to dampen gas flow into the nozzle and prevent particles of the gun propellant from entering the rocket motor. A rocket motor may implement the rigid structure or the baffled end cap, or both structures.

Abstract: A method of thrust vectoring a missile utilizing jet tabs is presented. Jet tabs are used to create lateral control moments on a missile by rotating tabs into the rocket exhaust plume and changing the thrust deflection angle. The method includes simultaneously rolling the missile during the thrust vector maneuver in order to reduce the maximum tab exposure to the rocket plume. The method enables aggressive pitchover maneuvers while reducing the risk of tab failure due to excessive exposure.

Abstract: An apparatus is provided for an aircraft propulsion system. This apparatus includes a thrust reverser cascade extending longitudinally between a cascade forward end and a cascade aft end. The thrust reverser cascade extends laterally between a cascade first side and a cascade second side. The thrust reverser cascade extends radially between a cascade inner face and a cascade outer face. The thrust reverser cascade includes a plurality of vanes and a wall. The vanes are arranged in a longitudinally extending array along the cascade inner face. The vanes include a first vane and a second vane. The wall is configured to block gas flow radially through the thrust reverser cascade between the first vane and the second vane.

Abstract: A system for multiple burns from a solid fuel rocket motor may extinguish rocket fuel after the rocket has been ignited. The motor may be extinguished via rapid decompression of the combustion chamber. The fuel may then be reignited by a suitable igniter, and the process of extinguishing and reigniting may be repeated, enabling multi-burn maneuvers. A decompressive extinguishing plug nozzle may extinguish solid rocket fuel after the rocket has been ignited and/or keep a rocket in a disarmed (zero thrust) state until the rocket is to be armed. The nozzle may include a plug that mostly impedes the opening of the nozzle and an outer cowl that is movable to rapidly decompress the combustion chamber. This rapid decompression extinguishes the solid rocket fuel. In some aspects, the fuel can be reignited and extinguished multiple times.

Abstract: The present disclosure discloses a propulsion method based on liquid carbon dioxide phase change and a propulsion device. The method includes the following steps of: accommodating carbon dioxide in a thermally insulated container in a liquid phase form; transiently heating to convert the carbon dioxide from a liquid phase to a gas phase; and jetting carbon dioxide gas after the phase change in a predetermined direction by a predetermined jet-out amount so as to obtain a propulsion force.

Abstract: A rocket booster has an annular shape, with a casing defining an annular space therewithin, and a solid rocket fuel in the annular spacing. The casing may itself at least in part define an annular gap that functions as a nozzle for the rocket booster, with protruding tabs on the casing aiding in maintaining a uniform height of the annular gap. The rocket booster may be mechanically coupled to an object protruding from the back of a fuselage of a flight vehicle, such as a missile. For example, the rocket booster may be placed around an aft turbojet nozzle of the flight vehicle. This allows the rocket booster to be used in situations where primary propulsion must be running both before and after (and perhaps during) the firing of the rocket booster.

Abstract: Power device based on alkali-water reaction, having a reaction chamber to carry out a chemical reaction between water and alkali element, having an exhaust nozzle to exhaust the reaction products generated in the reaction chamber, with a siphon tube connected to the exhaust nozzle, a nozzle shutter plate sealing the exhaust nozzle, and an external pressure inlet. The device has a water reservoir to store transfer water to the reaction chamber and an alkali reservoir to store and transfer an alkali element to the reaction chamber, and transfer device connecting the reaction chamber to both reservoirs to transmit the pressure generated in reaction chamber by part of the reaction products to the reservoirs, providing the transfer of water and alkali element from both reservoirs to the reaction chamber.

Abstract: A nozzle comprises a combustion chamber having a downstream end and a divergent formed of a cone-shaped wall extending between an upstream end and a downstream end. The upstream end of the divergent is connected to the downstream end of the combustion chamber by an intermediate ring having an upstream flange fixed on a fixing flange secured to the combustion chamber and a downstream flange connected to the upstream end of the divergent. The intermediate ring having an inner channel present between the upstream and downstream flanges of the intermediate ring. A material able to take heat from the ring is present in the inner channel.

Abstract: An injector for a liquid rocket engine includes an array of injector elements. Each injector element includes a central passage and a plurality of peripheral transverse passages. The central passages are configured to provide axial injection flow and the peripheral transverse passages are configured to provide swirl injection flow about the axial injection flow. A portion of the injector elements are configured to provide the swirl injection flow in a clockwise direction and another portion of the injector elements are configured to provide the swirl injection flow in a counter-clockwise direction. The injector elements are arranged to form a plurality of circumferential rows. The injector elements of each individual circumferential row are either all of the clockwise direction or all of the counter-clockwise direction. At least one of the circumferential rows is of the clockwise direction and at least one of the circumferential rows is of the counter-clockwise direction.

Abstract: A system and methods are disclosed for an upper stage space launch vehicle that uses gases from the propellant tanks to power an internal combustion engine that produces mechanical power for driving other components including a generator for generation of electrical current for operating compressors and fluid pumps and for charging batteries. These components and others comprise a thermodynamic system from which system enthalpy may be leveraged by extracting and moving heat to increase the efficient use of propellant and the longevity and performance of the launch vehicle.

Abstract: An integrated propulsion system for hybrid rockets includes an oxidizer tank, a rocket engine, a pressurization device, a pressurization device and an oxidizer pipe and valve unit. The rocket engine is disposed within the oxidizer tank partially and located on a first side of the oxidizer tank. The pressurization device is disposed, at least in part, within the oxidizer tank, is located on a second side of the oxidizer tank opposite to the first side of the oxidizer tank, and is configured to regulate an overall pressure level within the oxidizer tank. The oxidizer pipe and valve unit is connected to the oxidizer tank and the rocket engine, and is configured to control feeding of an oxidizer from the oxidizer tank into the rocket engine.

Abstract: A rocket configured to employ superheated water as a propellant, includes a photon reactor including a photon generator configured to generate UV light to superheat water.

Abstract: The inventors introduce the Borissov-Markusic Cycle as the new rocket engine cycle to solve the problem of low efficient open gas generator or tap-off gas generator cycles used to supply power to turbopump. A liquid rocket engine directs turbopump exhaust from a turbopump to a booster engine having an intake to accept ambient airflow, such as a variation of a ramjet, scramjet or dual mode ram scramjet engine. The turbopump is powered by combustion gases, such as from a gas generator or a tap-off manifold interfaced with the liquid rocket engine combustion chamber, and applies energy of the combustion gases to pump fuel and/or liquid oxygen to the liquid rocket engine combustion chamber. The combustion gases have a fuel-rich composition that includes unconsumed fuel from incomplete oxidation so that, upon injection into the combustion chamber of the booster engine, oxidation by ambient air of the unconsumed fuel releases energy to generate thrust with the booster engine.

Abstract: A liquid rocket engine assembly comprising a thrust chamber, a nozzle, and a joint structure. The joint structure attaches the thrust chamber and the nozzle and comprises at least one seal element and an attachment ring interposed between the thrust chamber and the nozzle. Fasteners extend between the nozzle and the thrust chamber through the at least one seal element and the attachment ring. Materials of the thrust chamber and of the nozzle comprise different coefficients of thermal expansion. A method of forming a liquid rocket engine assembly is also disclosed.

Abstract: A combustion chamber section for a combustion chamber for a rocket engine, the combustion chamber section including a combustion chamber body enclosing a combustion chamber volume and having coolant channels disposed therein, and at least one baffle integrally formed with the combustion chamber body and projecting from the combustion chamber body into the interior of the combustion chamber. The at least one baffle includes at least one coolant channel fluidly connected to at least one of the coolant channels in the combustion chamber body. Furthermore, an additive layer manufacturing method for manufacturing such a combustion chamber section is described.

Abstract: A power generation system includes a manifold having multiple plenums, capable of receiving multiple solid rocket motors. Initiators are coupled to the manifold, and are operatively coupled to respective of the plenums, to selectively fire different groups of the rocket motors coupled to respective of the plenums. The rocket motors act in parallel, to provide thrust in a single direction. The initiators may activate ignition charges that are in the plenums. The plenums may be annular plenums, which may be located in an annular manifold. The plenums may be lined with an insulator material. A cover may be used to cover the plenums, and also to receive the rocket motors. The rocket motors may be solid-fuel rocket motors, with propellant grains and nozzles. The individual rocket motors may have separate ignition booster charges coupled to the plenum, which are ignited by the ignition charge.

Abstract: A rocket booster has an annular shape, with a casing defining an annular space therewithin, and a solid rocket fuel in the annular spacing. The rocket booster also includes one or more nozzle pieces, mechanically coupled to the casing, that define one or more nozzles at the aft side of the rocket booster. The rocket booster may be mechanically coupled to an object protruding from the back of a fuselage of a flight vehicle, such as a missile. For example, the rocket booster may be placed around an aft turbojet nozzle of the flight vehicle. This allows the rocket booster to be used in situations where primary propulsion must be running both before and after (and perhaps during) the firing of the rocket booster.

Abstract: A hybrid metal composite (HMC) structure comprises a first tier comprising a first fiber composite material structure, a second tier longitudinally adjacent the first tier and comprising a first metallic structure and a second fiber composite material structure laterally adjacent the first metallic structure, a third tier longitudinally adjacent the second tier and comprising a third fiber composite material structure, and a fourth tier longitudinally adjacent the third tier and comprising a second metallic structure and a fourth fiber composite material structure laterally adjacent the second metallic structure. At least one lateral end of the second metallic structure is laterally offset from at least one lateral end of the first metallic structure most proximate thereto. Methods of forming an HMC structure, and related rocket motors and multi-stage rocket motor assemblies are also disclosed.